1. Field of the Invention
This invention relates to the heat treatment of aluminium-lithium alloys and in particular to such heat treatment for the strengthening of such alloys and for the optimisation of such alloys' plane stress fracture toughness. Such alloys are known in particular for use in aircraft skin construction, and more particularly for commercial aircraft fuselage, wing and empennage construction. In this application in particular the low density, high stiffness and excellent fatigue properties of aluminium-lithium alloys enable weight savings to be achieved to maximise profitability of the aircraft.
2. Discussion of Prior Art
Prior Art references which are relevant to this invention known at the time of the invention are as follows. "Effect of thermal exposure at 70.degree. C. on the performance of damage tolerant aluminium-lithium alloy sheet." February 1995. Reference DRA/SMC/WP952008 by D. S. McDarmaid; "Mechanical properties of 2024-T3 aluminium alloy sheet. December 1991. Reference TR91071 by D. S. McDarmaid, C. E. Thomas and C. Wheeler.
The aluminium-lithium (Al--Li) alloys registered with the ALUMINUM ASSOCIATION as AA8090 and AA2091 (hereinafter referred to without the "AA" prefix) in recrystallised sheet form and under-aged tempers have been shown to possess attributes of "Damage Tolerance" in that fatigue crack growth rates are commendably slow coupled with reasonably high levels of plane stress fracture toughness (Kc). A such, both products have been extensively investigated as potential alternatives to the currently most widely used materials for civil aircraft skin applications, in particular for fuselages eg alclad 2024 T3 and 2014A T4 sheet, where the density reduction associated with the lithium-bearing alloys would enable considerable amounts of weight to be saved. 8090 in plate form has also been investigated for upper and lower wing skin and empennage applications and may also be considered for upper wing skins.
In addition to the requirement for damage tolerance there are several other necessary characteristics which any new skin material and particularly fuselage wing and empennage skin materials must possess. These include adequate strength, good corrosion resistance and an often unstated but very important requirement of long-term thermal stability, ie the ability to withstand prolonged periods at moderately elevated temperatures without an appreciable or unacceptable loss in any of the key attributes. For a sub-sonic civil aircraft fuselage the worst case from a consideration of thermal instability involves on the ground exposures to the combined effects of high ambient temperatures and intense solar radiation. It is generally accepted that in tropical conditions fuselage skin temperatures of up to 70.degree.-85.degree. C. can be achieved when the sun is at or near its zenith. Over the life of an aircraft this could, in the worst case, represent a cumulative high temperature exposure of approximately 65000 hours (eg 6 hours per day for 30 years) although such an exposure would only be achieved for aircraft either stored in desert conditions or operated irregularly from tropical bases. Thermal stability is also one aspect of concern when considering the use of Al--Li alloys for wing and empennage skin applications.
The 8090 and 2091 alloys have been primarily investigated for fuselage skin applications in the T81 and T84 conditions respectively. The T81 condition for 8090 is achieved by artificial age hardening ("ageing") from the T31 condition (ie solution treated and controlled stretched) for 24 hours at 150.degree. C. whilst the T84 condition for 2091 is achieved by ageing from the T3 condition for 12 hours at 135.degree. C. following a slow ramp up from ambient to 135.degree. C. These treatments are intended to produce products which mimic the mechanical properties of alclad 2024 T3 (ie the lower limit for 0.2% Proof Stress has been set as approximately 270 MPa) in order that substitutional applications can more easily be considered. There is, also, the widespread perception that Al--Li alloys require static strengths at least equivalent to alclad 2024 T3 to be successful in the fuselage skin application. This is not necessarily so since the increase in Young's Modulus associated with the lithium content is capable of more than off-setting any slight reduction in strength which might now be seen to be required in order to properly satisfy a real requirement for very high fracture toughness and good impact resistance.
Despite the use of artificial ageing treatments, both the Al--Li products referred to are known to lack thermal stability in the temperature range 70.degree.-85.degree. C. and an increase in strength coupled with a disproportionately large reduction in Kc results after relatively short isothermal exposures (ie a very significant effect after 1000 hours). This inverse relationship between strength and Kc for Al--Li alloys has been demonstrated on many occasions. Given that the initial toughness levels for both alloys aged to their respective prior art conditions (ie T81 and T84 for 8090 and 2091 respectively) are marginal for the intended application when compared to alclad 2024 T3 (the current industry standard) this absence of thermal stability and the pernicious effect on toughness of even apparently very small increases in strength is widely regarded as a major contributory factor accounting for the current lack of any significant civil aircraft fuselage applications.
The cause of thermal instability is attributed to an on-going precipitation of .delta.' (Al.sub.3 Li). The reason for the continued precipitation of .delta.', and hence the thermal instability, is that there is an inverse relationship between the equilibrium volume fraction of .delta.' and temperature (ie the equilibrium volume fraction increases as temperature is reduced). The high rate of diffusion of lithium in aluminium ensures that the formation of .delta.' is not effectively diffusion rate controlled until the temperature falls some considerable way below the exposure temperature of concern. It therefore follows that even extensive ageing at the stated prior art ageing temperatures (ie 135.degree.-150.degree. C.) will never achieve anything approaching a complete precipitation of .delta.' and a high thermodynamic driving force for on-going precipitation, coupled with adequate rates of lithium diffusion, will exist at temperatures at or close to (below) the maximum thermal exposure temperatures considered. Instead, extensive ageing at these "higher" temperatures only serves to increase the volume fraction of other phases such as S' (Al.sub.2 CuMg) leaving a structure overly high in strength but relatively low in .delta.'. Subsequent long-term thermal exposure therefore results in a large increase in the .delta.' volume fraction, an increase in strength and embrittlement.
To illustrate the effect of on-going .delta.' precipitation duplicate samples of a batch of (hereinafter referred to as "Batch 1" material) 8090 T81 were given a range of thermal treatments prior to being exposed to an elevated temperature for a considerable length of time. The composition in weight percent of the Batch 1 material was:
______________________________________ Li Cu Mg Fe Zr Al ______________________________________ 2.23 1.14 0.79 0.045 0.06 Remainder ______________________________________
The treatments chosen included a 10 minute "reversion" at 200.degree. C. from the T81 condition (ie causing a drop in 0.2% Proof Stress due to .delta.' dissolution), followed by a re-age of 170.degree. C. for 4 hours (ie to achieve a recovery to approximately the original level of T81 0.2% Proof Stress and, finally, an extensive over-ageing treatment of 220.degree. C. for 12 hours in addition to the T81 initial treatment.
After tensile testing one long transverse (LT) oriented sample representative of each condition the duplicate samples of all conditions including the T81 "Control" condition were then exposed for 920 hours at 100.degree. C. in order to crudely represent a lifetime's exposure to tropical temperatures. The results of the mechanical property tests and electrical conductivity measurements made are shown in Table 1.
It is clear from Table 1 that the on-going precipitation at 100.degree. C. results in a considerable increase in strength. The reverted material recovers to a higher strength than is the case for the Control condition indicating the ineffectiveness of reversion as a means of increasing the toughness of 8090 where consideration must also be made of thermal instability effects since the initial benefit of reversion is short-lived and the treatment can, ultimately, be expected to be harmful as it results in a higher final strength after thermal exposure. The increase in strength of the reverted material over and above the un-reverted material at the conclusion of the thermal exposure is attributed to the additional S' precipitated during the reversion process. Similarly, the additional increase in strength of the reverted and re-aged material following thermal exposure compared with either of the T81 and T84 plus reversion conditions is attributed to the increased S' associated with 4 hours at 170.degree. C.
Finally, the use of over-ageing is seen to be completely ineffective at achieving stability with a 48 MPa rise in 0.2% Proof Stress being apparent at the conclusion of the 920 hour exposure. Similar results for all starting conditions would be anticipated for exposure at, say, 70.degree. C. and an even higher equilibrium volume fraction of .delta.' would be realisable at this temperature than at 100.degree. C. although the exposure time required to achieve saturation would be that much greater at the lower temperature due to the reduced diffusion rates.
It should be noted that the Batch 1 8090 sheet had a T81 LT 0.2% Proof Stress of 293 MPa and which then reached what is believed to be a .delta.' -saturated 0.2% Proof Stress of 320 MPa following 920 hours thermal exposure at 100.degree. C., ie a rise of 27 MPa.